Retractable z-fold flexible blanket solar array

ABSTRACT

A solar array structure for a spacecraft includes one or a pair of flexible blanket or other foldable solar arrays and a deployable frame structure. The deployable frame structure includes a T-shaped yoke structure, a T-shaped end structure, and one or more rigid beams, the T-shaped yoke structure connectable to the spacecraft. When deployed, the frame structure tensions the flexible blanket solar array or arrays between the T-shaped yoke structure and the T-shaped end structure. When stowed, the flexible blanket solar array or arrays are folded in an accordion manner to form a stowed pack or packs between the cross-member arms of the T-shaped yoke structure and the T-shaped end structure, also stowed in its own Z-fold arrangement. The cross-member arms of the T-shaped end structure can include a solar array that can provide power before deployment while the flexible blanket solar array is stowed.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a Continuation-in-Part of U.S. patent applicationSer. No. 17/398,319, filed Aug. 8, 2021, which is incorporated herein byreference.

BACKGROUND

To provide operating power, satellites use solar array structures with alarge surface area of photovoltaic cells to generate electricity fromthe sunlight incident on the array structure. For shipment and launchthe solar array is stowed to have a small volume and then deployed oncethe spacecraft has been launched. For launch purposes, the smaller thevolume and the lower the weight, the better. Once fully deployed, it isdesirable that the solar array structure provide a light weight, stiff,strong, stable, and flat surface of sufficient surface area that canallow uniform exposure to the sun and minimize on-orbit spacecraftattitude control disturbance while meeting the satellite's powerrequirements. These conflicting needs result in an ongoing pursuit ofimprovements in the design of such solar arrays.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a block diagram of a spacecraft system.

FIG. 2 is a block diagram of an example spacecraft.

FIGS. 3 and 4 illustrate two views of a spacecraft with deployed solararrays.

FIGS. 5A, 5B, and 5C illustrate a first embodiment of a Z-fold flexibleblanket solar array in a stowed configuration, partially deployed, and adeployed configuration, respectively.

FIG. 5D is a side view of the deployed Z-fold flexible blanket solararray structure of FIGS. 5A, 5B, and 5C.

FIGS. 6A and 6B illustrate another embodiment of a Z-fold flexibleblanket solar array in a stowed configuration and a deployedconfiguration, respectively.

FIGS. 7 and 8 are side views of a stowed Z-fold flexible blanket solararray with two different orientations relative to the spacecraft.

FIGS. 9A-9D illustrate an embodiment in which the solar panels on the onthe arms of the T-shaped support structures fold when stored.

FIGS. 10 and 11 are flowcharts respectively describing embodiments forthe stowing and the deploying of the Z-fold flexible blanket solar arraystructures.

FIG. 12 is a side view of an embodiment for a motorized large angleflexure pivot hinge that can be used in the frame structure for z-foldflexible blanket solar array structures.

FIG. 13 is an oblique view of the motorized flexure hinge showing thefeatures of FIG. 12 and also the inner rim.

FIGS. 14A and 14B respectively illustrate a bimetallic latch arm in areleased state and a latched state.

FIGS. 15 and 16 illustrate an embodiment for a deployed z-fold flexibleblanket solar array structure in which the yokes and intermediatemembers are at less than 180° when latched in a deployed configuration.

FIG. 17 illustrates an embodiment for the locating of the controller andwiring for the motorized flexure hinges and releasable latches.

FIG. 18 is a flowchart for an embodiment of a method of operating aretractable Z-fold flexible blanket solar array structure.

DETAILED DESCRIPTION

A solar array structure for a satellite or other spacecraft is made upof one or a pair of flexible blanket solar arrays and a deployable framestructure. The deployable frame structure includes a T-shaped yokestructure, a T-shaped end structure, and one or more rigid beams, wherethe T-shaped yoke structure connects to the spacecraft. When deployed,the frame structure tensions the flexible blanket solar array or arraysbetween the T-shaped yoke structure and the T-shaped end structure. Whenstowed, the flexible blanket solar array or arrays are folded in anaccordion manner to form a stowed pack or packs between the cross-memberarms of the T-shaped yoke structure and the T-shaped end structure ofthe frame structure, which is also stowed in its own Z-fold arrangement.The surface of the cross-member arms of the T-shaped end structure caninclude a solar array that can provide power before deployment while theflexible blanket solar array is still stowed. To allow for longercross-arms, and a larger area exposed solar array when stowed, in someembodiments the cross-member arms can fold inward when stowed.

FIG. 1 is a block diagram of a spacecraft system that can implement thetechnology proposed herein. The system of FIG. 1 includes spacecraft 10,subscriber terminal 12, gateway 14, and ground control terminal 30.Subscriber terminal 12, gateway 14, and ground control terminal 30 areexamples of ground terminals. In one embodiment, spacecraft 10 is asatellite; however, spacecraft 10 can be other types of spacecrafts(e.g., shuttle, space station, inter-planet traveling craft, rocket,etc.). Spacecraft 10 may be located, for example, at a geostationary ornon-geostationary orbital location. Spacecraft 10 can also be a LowEarth Orbit satellite. Spacecraft 10 is communicatively coupled by atleast one wireless feeder link to at least one gateway terminal 12 andby at least one wireless user link to a plurality of subscriberterminals (e.g., subscriber terminal 12) via an antenna system. Gatewayterminal 14 is connected to the Internet 20. The system allowsspacecraft 10 to provide internet connectivity to a plurality ofsubscriber terminals (e.g., subscriber terminal 12) via gateway 14.Ground control terminal 30 is used to monitor and control operations ofspacecraft 10. Spacecraft can vary greatly in size, structure, usage,and power requirements, but when reference is made to a specificembodiment for the spacecraft 10, the example of a communicationsatellite will often be used in the following, although the techniquesare more widely applicable, including other or additional payloads suchas for an optical satellite.

FIG. 2 is a block diagram of one embodiment of spacecraft 10, which inone example (as discussed above) is a satellite. In one embodiment,spacecraft 10 includes a bus 202 and a payload 204 carried by bus 202.Some embodiments of spacecraft 10 may include more than one payload. Thepayload provides the functionality of communication, sensors and/orprocessing systems needed for the mission of spacecraft 10.

In general, bus 202 is the spacecraft that houses and carries thepayload 204, such as the components for operation as a communicationsatellite. The bus 202 includes a number of different functionalsub-systems or modules, some examples of which are shown. Each of thefunctional sub-systems typically include electrical systems, as well asmechanical components (e.g., servos, actuators) controlled by theelectrical systems. These include a command and data handling sub-system(C&DH) 210, attitude control systems 212, mission communication systems214, power subsystems 216, gimbal control electronics 218 that be takento include a solar array drive assembly, a propulsion system 220 (e.g.,thrusters), propellant 222 to fuel some embodiments of propulsion system220, and thermal control subsystem 224, all of which are connected by aninternal communication network 240, which can be an electrical bus (a“flight harness”) or other means for electronic, optical or RFcommunication when spacecraft is in operation. Also represented are anantenna 243, that is one of one or more antennae used by the missioncommunication systems 214 for exchanging communications for operating ofthe spacecraft with ground terminals, and a payload antenna 217, that isone of one or more antennae used by the payload 204 for exchangingcommunications with ground terminals, such as the antennae used by acommunication satellite embodiment. The spacecraft can also include anumber of test sensors 221, such as accelerometers that can used whenperforming test operations on the spacecraft. Other equipment can alsobe included.

The command and data handling module 210 includes any processing unit orunits for handling includes command control functions for spacecraft 10,such as for attitude control functionality and orbit controlfunctionality. The attitude control systems 212 can include devicesincluding torque rods, wheel drive electronics, and control momentumgyro control electronics, for example, that are used to monitor andcontrol the attitude of the space craft. Mission communication systems214 includes wireless communication and processing equipment forreceiving telemetry data/commands, other commands from the groundcontrol terminal 30 to the spacecraft and ranging to operate thespacecraft. Processing capability within the command and data handlingmodule 210 is used to control and operate spacecraft 10. An operator onthe ground can control spacecraft 10 by sending commands via groundcontrol terminal 30 to mission communication systems 214 to be executedby processors within command and data handling module 210. In oneembodiment, command and data handling module 210 and missioncommunication system 214 are in communication with payload 204. In someexample implementations, bus 202 includes one or more antennae asindicated at 243 connected to mission communication system 214 forwirelessly communicating between ground control terminal 30 and missioncommunication system 214. Power subsystems 216 can include one or moresolar panels and charge storage (e.g., one or more batteries) used toprovide power to spacecraft 10. Propulsion system 220 (e.g., thrusters)is used for changing the position or orientation of spacecraft 10 whilein space to move into orbit, to change orbit or to move to a differentlocation in space. The gimbal control electronics 218 can be used tomove and align the antennae, solar panels, and other external extensionsof the spacecraft 10.

In one embodiment, the payload 204 is for a communication satellite andincludes an antenna system (represented by the antenna 217) thatprovides a set of one or more beams (e.g., spot beams) comprising a beampattern used to receive wireless signals from ground stations and/orother spacecraft, and to send wireless signals to ground stations and/orother spacecraft. In some implementations, mission communication system214 acts as an interface that uses the antennae of payload 204 towirelessly communicate with ground control terminal 30. In otherembodiments, the payload could alternately or additionally include anoptical payload, such as one or more telescopes or imaging systems alongwith their control systems, which can also include RF communications toprovide uplink/downlink capabilities.

FIGS. 3 and 4 look an exterior view for an embodiment of spacecraft 10in more detail. More specifically, FIGS. 3 and 4 show two views of anembodiment of spacecraft 10, where FIG. 4 shows the spacecraft rotatedby 90° about the axis of the solar arrays 265 relative to FIG. 3 . Anumber of different embodiments are possible, but the example of FIGS. 3and 4 can be used to illustrate some of the elements relevant to thecurrent discussion.

Referring to FIGS. 3 and 4 , the spacecraft 10 includes a spacecraftbody 261 from which extend two, in this example, deployed solar arrays265. Attached to the body will also be one or more number of antennae217 and 243 as described above, by which the satellite can receive andtransmit signals. Depending on the particulars of the embodiment, asatellite may have a large number of antennae, but only a pair ofantennae for exchanging signals with a ground station are shown.Attached to the spacecraft body 261 are a number of thrusters, as shownat 263 and 267, which typically include one or more main thrusters and anumber of attitude and orbit control thrusters. Internal to the bodywill be the spacecraft's frame (not show) within which the functionalsub-systems can be installed.

The deployed arrays 265 can include a solar array, a thermal radiatingarray, or both and include one or more respectively coplanar panels. Thedeployed arrays 265 can be rotatable by the gimbal control or solararray drive assembly 251 about the longitudinal axis (the left-rightaxis in FIGS. 3 and 4 ), in order to achieve or maintain a desiredattitude with respect to, for example, the sun. For embodiments in whichthe deployed arrays 265 include a solar array, the solar array may bearticulable so as to be substantially sun facing and electricallyconnected to the spacecraft 10 to provide power. The deployed solararray 265 may be sized and positioned so as to generate substantiallymore power from sunlight than would be possible if the solar array wasfixedly disposed on the body 261 of the spacecraft 10. For example, insome implementations, the solar array orientation may be rotatable aboutthe longitudinal axis of the spacecraft 10 so that photovoltaic powergenerating surfaces of the solar array remains substantially sun facing.

For shipping and launching of the spacecraft 10, the solar array isstowed into a small volume. Although the stowed volume is wanted to beas small as practicable, the solar array will also need to be largeenough to provide sufficient power for spacecraft operations oncedeployed. One set of embodiments for a solar array are based on aflexible photovoltaic blanket. Flexible blanket solar arrays use a thinflexible substrate that has mounted to it on one side an array ofphotovoltaic solar cells (and refractive concentrators) and associatedwiring that can be folded into a small and compact package for stowage;and is attached to the deployable solar array structure for unfurlingduring deployment into a flat, aligned, and tensioned configuration whenfully deployed. The following presents embodiments for a blanket solararray that can be compactly stowed and then efficiently deployed toprovide an effective solar array for spacecraft operations.

Once launched, a spacecraft will typically need to perform orbit raisingand other maneuvers that require power before the solar arrays can bedeployed and activated, where this power will usually need to beprovided by batteries within the power subsystems 216. As batteries addweight, they are often kept to near the minimum that can still meet thespacecrafts needs. Therefore, it is critical to deploy the solar arrayas soon as possible before the batteries are drained. In addition tobattery power, having a source of solar power before solar arraydeployment would be extremely beneficial to orbit raising operations ofa spacecraft as this could provide additional power and chargingcapability in contingency situations.

Although flexible blanket solar arrays of different shape and forms havebeen used for spacecraft previously, they have not been capable ofproviding power in their stowed configuration. Embodiments describedbelow present structural configurations that not only accommodates useof a blanket solar array, but also provide some power in its stowedconfiguration. As described in more detail in the following, this isaccomplished by stowing the blanket between two T-shaped rigidstructural members connected to each other by narrow structural beams.Stowed power is provided by solar cells mounted on the outboardT-structure. In one set of embodiments, the blanket solar array isstretched between the two T-shaped structures when deployed, and foldedbetween the two top wider sections of the T structures in its stowedconfiguration. Deployment of the structural members can be synchronizedby cable loops such that all members deploy in a predictable fashion andlock up to form a rigid structure with required strength and stiffness,where in addition to providing structure for the blanket solar array,the deployment structure provides a surface for mounting solar cells togenerate power in the stowed configuration.

More specifically, embodiments are presented here for a Z-fold flexibleblanket solar array structure. The structure includes a T-shape rigidstructure yoke, a T-shape end structure and sections of rigid beam,where the rigid structures are connected to each other through hingesand synchronized to deploy from the stowed to the deployedconfiguration. The blanket solar panel can be folded and stowed betweenthe inboard T-yoke and the outboard T-structure in stowed configuration,with deployment of rigid structure unfolding the blanket from stowed todeployed configuration. Embodiments can be configured where blanketsolar array is split in two sections, one on either side of thedeploying structure, or with a one piece blanket solar array, where thedeploying structure is located behind the blanket. The outboard surfaceof the T structure can be populated with solar cells to provide limitedpower in stowed configuration, and, in additional embodiments, thisconcept provides the option to double the array capability by stowingtwo arrays side-by-side on a single yoke.

FIGS. 5A, 5B, and 5C illustrate a first embodiment of a Z-fold flexibleblanket solar array 501 in a stowed configuration, partially deployed,and a deployed configuration, respectively. FIG. 5D is a side view ofthe deployed Z-fold flexible blanket solar array structure 501. FIGS. 5Aand 5B include the spacecraft 10 to which the solar array 501 isattached, while FIGS. 5C and 5C just show the array structure 501.

The Z-fold flexible blanket solar array structure 501 of FIGS. 5A-5Dincludes frame structure of a T-shaped rigid structure yoke 511,sections of rigid beam 513, and an outboard T-structure. The componentsof the rigid frame structure are connected to each other through thehinges 515 and are synchronized to deploy from the stowed configurationof FIG. 5A to the deployed configuration of FIG. 5C. The blanket solararray 521 is connected in its long direction at one end to the inboardT-yoke 511 and at the other end to the outboard T-yoke 517, butunconnected to the intermediate members 513. The blanket solar array 521is folded and stowed between the inboard T-yoke 511 and the outboardT-yoke 517 in its stowed configuration and can be held in place by oneor more hold-downs 519. The deployment of the rigid structure ofelements 511, 513, and 517 for its Z-fold stowed configuration can beimplemented through the hinges 515, which can be spring loaded, tounfold the blanket solar array 521 from the stowed configuration of FIG.5A, through the intermediate state of FIG. 5B, to the deployedconfiguration of FIG. 5C. The outboard T-structure 517 can be populatedwith a solar cell array 537 to provide stowed power and the inboardT-structure 511 can also be provided with a solar cell array 531.

Considering the deployed structure of FIG. 5C further, this is for aview where the cell side of the Z-fold flexible blanket solar arraystructure 501 is shown, with the blanket side underneath on the back.(Relative to FIGS. 5A and 5B, those two figures illustrate the stowedand the partially deployed structure where the blanket side is on theupper surface and the cell side on the lower surface.) In thisembodiment the blanket solar array has two parts, 521A and 521B, locatedto either side of the central rigid structure of the T-structures 511and 517 and rigid beams 513. In one set of embodiment, the backing ofthe blanket sections 521A and 521B can be formed of a mylar film.Embodiments for the rigid T-structures 511 and 517 and the rigid beams513 can, for example, be made of hollow graphite rectangular tubes withdimensions of a few inches and a wall thickness of 10s of mils. Thelength of the rigid T-structures 511 and 517 and the rigid beams 513 islonger than multiple ones of the of the foldable sections of the blanketsections 521A and 521B. The length of the rigid T-structures 511 and 517and the rigid beams 513, and the number of the rigid beams 513, dependson the chosen size of the blanket array when deployed and size allowablefor these pieces of the rigid structure when stowed, as discussed inmore detail with respect to FIGS. 7 and 8 .

Although this description is given in the context of a flexible blanketsolar array, other embodiments based on the use of flexible, semi-rigid,or rigid panels as substitute surfaces for the laydown and bonding ofphotovoltaic solar cells. As can be seen from the view of FIG. 5A andsimilar figures, if the panels are connected together by use of hingesor are other otherwise foldable so that they are foldable in thedirection extending (when deployed) away from the spacecraft (the axialdirection) to form a stowed pack of a plurality of folds, the techniquesand structures described here can be applied to such foldable solararrays. To give some examples, alternate embodiments of a foldable solararray could include a Kapton film, a flexible sheet of graphite a fewmils thick, or rigid or semi-rigid honeycomb panels. Although thefollowing will continue to refer to flexible blanket embodiments, itwill be understood that alternate embodiments can be based on otherfoldable solar arrays.

When deployed, the solar array blanket sections 521A and 521B areclamped in place between the T-shaped yoke structure 511 and theoutboard T-shaped structure 517. Each of the solar array blanketsections 521A and 521B can include a blanket guide line or string, suchas 541, that can extend from a retractable spool, for example, to helptension the solar array blanket sections 521A and 521B. The powergenerating capacity of the structure when deployed can be increased bythe solar arrays 531 and 537 on the structures 511 and 517, with thesolar array 537 on the outboard T-structure also supplying power whenstowed. To deploy and hold the structure in the extended structureillustrated in FIG. 5C, simple hinges and a closed cable loop, such asmade up of the elements 543, can be used to synchronize the deploymentprocess of the rigid T-structures 511 and 517 and the rigid beams 513,where the cable of the closed cable loop structure can be located to theinside of the structure avoid contact with the blanket section or thesolar cells.

Returning to FIG. 5A, the stowed configuration for the Z-fold flexibleblanket solar array (or other foldable array) structure 501 is shown.The Z-fold structure of the rigid T-structures 511 and 517 and the rigidbeams 513 is collapsed and held in place one or more hold-downs 519.Note that the solar cell array 537 of the outboard T-structure 517 isfacing outward and can, consequently, provide power even when thestructure is stowed. FIG. 5A only shows a single hold-down 519 aroundthe bottom of the stowed structure, but other hold-downs could also beplaced around the cross-member arms of the rigid T-structures 511 and517 or other additional locations, such as is shown in FIGS. 7 and 8below. The solar array blanket sections 521A and 521B are also Z-foldedin an accordion manner as a set of pleats to form a pack of as shown at521. The rigid beams 513 extend up between the blanket packs for solararray blanket sections 521A and 521B, so that in the view of FIG. 5Athese are seen to overlay the rigid beams 513. Deployment of the solararray structure 501 initiates upon release of the hold-downs 519. Notethat in FIG. 5A, as well as FIG. 5B, the blanket's back side (i.e., theside without the photovoltaic cells) is facing upward and the solar cellside is facing downward.

FIG. 5B illustrates the Z-fold flexible blanket solar array structure501 at an intermediate point during deployment. The hinges 515 caninclude a graphite or other spring structure, such as represented at thesides of the hinge 515, to provide a spring driven synchronizeddeployment of the structure to extent the blanket to the flat positionof FIG. 5C. The structural hinge-lines can be latched mechanically ormagnetically before tensioning the blanket sections 521A and 521B, withthe blanket tension then maintained by blanket spring tensioning.

FIG. 5D shows a side view of the deployed structure 501 where thesections of the solar array blanket 521 are oriented, as in FIGS. 5A and5B, with the photovoltaic cells on the bottom side and the blanketbacking to the top. The solar cell arrays 531 and 537 of the inboard andoutboard T-structures 511 and 517 are also facing downward. The rigidbeams 513 are connected by the hinges 515 and run between the solararray blanket sections 521A and 521B.

For the embodiment of FIGS. 5A-5D, as well as the other embodimentsdiscussed below, to control the speed a viscous or other damper can belocated where the T-shaped yoke structure 511 connects to thespacecraft's solar array drive assembly (such as incorporated into themount 551), which the spacecraft 10 can use to rotate the Z-foldflexible blanket solar array structure 501 to face the sun. In one setof embodiments, the hinges 515 can be simple hinges, of the pin type,without need of mono-ball bearings and where the structure does notrequire offset springs by use of center mounting. In other embodiments,mono-ball bearings, offset springs, or both can be used. Simple latchesor catches can be used to maintain the deployed position.

FIGS. 6A and 6B respectively illustrate the stowed and deployed Z-foldflexible blanket solar array (or other foldable array) structure 601 inan alternate embodiment having a one piece solar array blanket 621 andwhere the deploying structure is located behind (i.e., on the backsideof) the solar array blanket 621 when deployed. Relative to FIGS. 5A and5C, corresponding elements are similarly number: e.g., T-shaped yokestructure 511 is now 611, rigid beams 513 are now 613, and so on.

In the deployed view of FIG. 6B, the single piece solar array blanket621 is again shown with the photovoltaic cells facing outward toward theviewer, but the support structure of including the rigid beams 613 arenow on blanket side behind the solar array blanket 621. The outboardT-structure 617 can again be populated with a solar cell array 637 toprovide stowed power and the inboard T-structure 611 can also beprovided with a solar cell array 631. The closed cable loop can beinside of the deployment structure of the yoke T-structure 611, rigidbeams 613, and the outboard T-structure 617, so as to not pinch thesolar array blanket.

When stowed in a Z-fold, as shown in FIG. 6A, it can be seen that,relative to the embodiment of FIG. 5A, the rigid beams 613 arerelatively shorter so that, when stowed, they do not extend upward tothe blanket pack into which the solar array blanket 621 has been foldedfor stowing. The rigid beam 613 closest to the yoke T-structure 611 isnow also attached below the stowed solar array blanket 621, so that whendeployed the rigid beam 613 closest to the yoke T-structure 611 willpartially lie below the yoke T-structure 611.

FIGS. 7 and 8 illustrates a side view of a spacecraft 10 with the Z-foldflexible blanket solar array or other foldable array in a stowedconfiguration. A spacecraft 10, such as a satellite will often be of arectangular shape on its sides, such as being taller than wide in theexamples of FIGS. 7 and 8 . The stowed Z-fold flexible blanket solararray in a stowed configuration can be attached to the spacecraft withtwo orientations. Of the Z-fold flexible blanket solar array, theoutboard T-yoke 717/817, including the solar cell array 737/837, can beseen in this view, with the rest of structure stowed behind. Threehold-downs 719/819 are shown in these embodiments, including a hold-downon the lower part of the base of the yoke T-structure 717/817 arrangedas for the hold-down 517/617 of FIG. 5A/6A, and a hold-down on each ofthe arms.

The embodiments of FIGS. 7 and 8 have different relative advantages andthe choice is a design decision based on the requirements of aparticular spacecraft. The embodiment of FIG. 7 allows for a longeroutboard T-yoke 717 and longer rigid beams (not visible in the view ofFIG. 7 ), allowing for a greater extension when deployed for a givennumber of rigid beams. FIG. 8 allows for a wider longer outboard T-yoke717 and, consequently, a wider flexible blanket solar array. Theembodiment of FIG. 8 also allows for a larger solar cell array 837, andhence more power when stowed, than the solar cell array 737 of FIG. 7 .It should be noted that in FIGS. 7 and 8 , and the similar FIGS. 9A-9D,the elements may not be represented to scale relative to one another.For example, the size of the spacecraft may be several feet in eitherdirection, while the width of the T-yoke structure only a few incheswide.

FIGS. 9A-9D present another embodiment for a Z-fold flexible blanketsolar array that allows for a larger usable solar array when stowed. Theembodiment of FIGS. 9A-9D can provide the spacecraft with roughly doublethe pre-deployment power capability by stowing two arrays side-by-sideon a single yoke. The configuration of this embodiment uses a 90°rotation of the stowed blanket stacks to be perpendicular to the centralmember of the yoke and end structures prior to deployment.

FIGS. 9A-9C respectively illustrate the Z-fold flexible blanket solararray with foldable arms in its initial stowed position, with the armspartially rotated, and fully rotated to form a T-structure. The view ofFIG. 9A is the same as for FIG. 7 or 8 , but hold-downs are not shown tosimplify the presentation. Relative to FIG. 7 , rather than for a fixedT-shaped structure, the top member of T-shaped yoke now includes twofoldable sections covered with outward facing solar arrays 937A and 937Bthat rotate by use of respective hinges 947A and 947B. This provides foran increased available surface array for the solar arrays 937A and 937B,and consequently more available power, while still allowing for thestructure to be stowed within the form-factor of the spacecraft 10. Oncedeployed, the resultant Z-fold flexible blanket solar array will also bewider, again resulting in more available power for a given length ofstructure. The two curved arrows FIGS. 9A and 9B are not part of thestructure, but meant to indicate how the arms will rotate in an initialphase of the deployment process.

As for the hold downs in the stowed configuration of FIG. 9A, ahold-down could again be placed on lower part (as seen in this review)of the central member 917 of the yoke section and a hold-down on thelower part of each of the folded arms with the outward facing solararrays 937A and 937B. One or more additional hold-downs can also on theupper part of one or more of these elements. In one embodiment, the armsare first folded out, after which the structure can then be furtherdeployed similarly to the embodiment of FIGS. 5A-5D. FIGS. 9A-9Cillustrate this initial deployment phase.

Starting from FIG. 9A, the hold-downs for the arms 937A and 937B arereleased. Similar to the hinges 515 as discussed above, the hinges 947Aand 947B can include a spring structure that will then cause the arms937A and 937B to rotate outward at the bottom in the direction of thecurved arrows. FIG. 9B shows an intermediate position where the arms937A and 937B have partially rotated. In FIG. 9C the arms 937A and 937Bhave finished rotating and can be latched in place to the central member917 of the yoke to complete the initial deployment phase and form aT-structure similar to that shown in FIG. 7 , but with the top sectionof the T now extending beyond sides of the spacecraft 10. At this point,when viewed from the side (i.e., rotating the view by 90° about thevertical axis in the positive direction), the structure would appear asin FIG. 5A, although extending further into and out of the page. Oncethe side arms are latched in place by a mechanical or magnetic latch,for example, the structure can then complete the deployment process muchas illustrated by FIG. 5B, where the resultant deployed Z-fold flexibleblanket solar array 901 is shown in FIG. 9D.

FIG. 9D shows the deployed configuration of an embodiment of the Z-foldflexible blanket solar array 901. The view of FIG. 9D is similar to thatof FIG. 5D and the components are similarly number: i.e., rigid beams513 are now 913, lower solar array blanket 521A is now 921A, and so on.Aside from the inboard and outboard T-structures, the Z-fold flexibleblanket solar array 901 can be as described above with respect to theZ-fold flexible blanket solar array 501 of FIG. 5D. The outboardT-structure is now formed of a central member 917 and the two foldingarms covered with solar arrays 937A and 937B and connected by the hinges947A and 947B. Although not visible in the views of FIGS. 9A-9C, theyoke of the inboard T structure is similarly formed of a central member911 and the two folding arms that can be covered with solar arrays 937Aand 937B and connected by the hinges 947A and 947B, so that when stowedthe arms of both the inboard and outboard T-structures are foldeddownward with the folded blanket packs for solar array blankets 921A and921B stowed between them.

As with the embodiments of FIGS. 5A-5D and FIGS. 6A and 6B, theembodiment+ of FIGS. 9A-9D can again use a viscous or other damper tocontrol the deployment speed can be located where the T-shaped yokestructure's central member 911 connects to the spacecraft's solar arraydrive assembly and embodiments for the hinges (both 945A/B and 947A/B,as well as the equivalent of 515) can again be simple hinges, of the pintype, without need of mono-ball bearings in a structure does not requireoffset springs by use of center mounting. In other embodiments,mono-ball bearing, offset springs, or both can be used for these hinges.Simple latches or catches can be used to maintain the deployed position.To deploy and keep in place the Z-fold flexible blanket solar array 901,a closed cable loop, that can include a concealed cable, can be usedalong with simple latches or catches without rollers. In someembodiments, the Z-fold flexible blanket solar array 901, as well as theother embodiments, can be motorized to be retractable.

Relative to the embodiments of FIGS. 5A-5D and FIGS. 6A and 6B, the useof the folded arms allows for a larger exposed surface area so that, bypopulating these with solar cells to form the arrays 937A and 937B, morepower is available to the spacecraft when the blanket solar array 901structure is stowed. The longer arms of the inboard and outputT-structure also allow for wider solar array blankets 921A and 921B and,if populated with cells, longer solar arrays 931A/B and 937A/B. In someembodiments, one or both of the central members 911 and 917 of theT-structures can also be populated with solar cells.

FIGS. 10 and 11 are flowcharts respectively describing embodiments forthe stowing and the deploying of the Z-fold flexible blanket solar arraystructures. FIG. 10 begins at step 1001 with receiving, which caninclude the complete or partial construction, of the frame of thedeployment structure including the T-shaped yoke structure(511/611/911), the end T-shaped structure (517/617/917), the one or morerigid members (513/613/913), and the hinges (515/615 and 945A/B, 947A/Bin the embodiment of FIGS. 9A-9D). At step 1003 the flexible blanketsolar array is received, which can include the complete or partialconstruction of the flexible blanket solar array. Step 1003 can beperformed before, after, or concurrently with step 1001.

The flexible blanket solar array is attached to the frame of thedeployment structure at step 1005, with one end of the blanket orblankets being attached to the yoke cross-member arm of T-structure(511/611/911) and other end to the cross-member arm of the endT-structure (517/617/917). The blanket solar array can already befolded, or partially folded, in an accordion manner along its length toform a pack when attached, or can be folded after attachment, such aspart of step 1007. The full structure (501/601/901) is then placed intoits stowed configuration at step 1007 by folding up the frame of thedeployment structure and, if not already done, the blanket solar arrayso that the blanket pack or packs are between the cross-arms of theT-structures (511/611/911) and (517/617/917). In an embodiment such asthat of FIGS. 9A-9D, the arms of the T-structures would also be foldeddown into the arrangement shown in FIG. 9A.

In step 1009 the Z-fold flexible blanket solar array structure(501/601/901) is attached to the spacecraft 10 by bottom end of thecentral member of the yoke T-structure (511/611/911) as shown in theside views of FIG. 5A or 6A, for example. Depending on the embodiment,the flexible blanket solar array structure (501/601/901) can be attachedafter being placed into a stowed configuration at step 1007, or can beassembled or partially assembled (parts or all of steps 1001, 1005, or1005) while attached and then placed into the stowed figuration, so thatstep 1009 could be performed at some point before step 1007. Onceattached and in stowed configuration, releasable hold-downs(519/619/719/819) can be attached at step 1011 to hold the Z-foldflexible blanket solar array structure (501/601/901) in place untildeployment.

FIG. 11 is a flowchart of an embodiment of the deployment process andbegins once the spacecraft has been launched. Starting at step 1101, forembodiments that include photovoltaic solar cell arrays(527/637/737/837/937A,B) on the cross-arms of the outboard T-structure(517/617/717/817/917), the blanket solar array structure (501/601/901)can provide power prior to deployment. At step 1103, the controlcircuits on the spacecraft, such as the command and data handlingsub-system (C&DH) 210, initiates a deployment command. This command canbe received from a ground station or originate on the spacecraft itself,such as in response to a specified set of conditions.

For embodiments such as in FIGS. 9A-9D in which the horizontal arms folddown for stowing, as shown for the solar array populated arms 937A,B asshown in FIG. 9A, an initial deployment phase 1104 can be included. Atstep 1105 hold-downs for the arms are released, such as by pulling outof a pin by a servo or actuator. Once released, the solar arraypopulated arms 937A,B begin to rotate about the hinges 947A,B from theposition of FIG. 9A, through the position of 9B, and into the positionof 9C at step 1107. The arms 931A,B will also rotate about the hinges945A,B at the same time, although these are not visible in FIGS. 9A-9Cas they are behind the end structure arms in these views, with theblanket packs stowed between them. The rotation can be effected by useof graphite or other spring structures incorporated into the structureof the hinges 945A,B and 947A,B or other automatic device or couldincorporate a motor. The arms are then latched into place at step 1109,such as by mechanical or magnetic latches, where these could beincorporated into the hinges 945A,B and 947A,B, separate latches, or acombination of these.

At step 1111 the hold-downs (519/619/719/819) are released by thecontrol circuits of the spacecraft, such as by pulling out of a pin by aservo or actuator, and the Z-fold flexible solar array blanket and framebegin to expand at step 1113, such as is illustrated in FIG. 5B. Theexpansion can be implemented by use of graphite or other springstructures incorporated into the structure of the hinges (515/615) orother automatic device or could incorporate a motor. As discussed above,simple hinges and a closed cable loop can be used to synchronize thedeployment process and the speed can be controlled by used of a viscousor other damper can be located where the T-shaped yoke structure(511/611/911) connects to the spacecraft's solar array drive assembly.The guide line or string (541/641/911), that can extend from aretractable spool, for example, can be used to help tension the solararray blanket sections, both once deployed and also during step 1113.The structure is then latched into its deployed arrangement at 1115,such as by mechanical or magnetic latches, where these could beincorporated into the hinges (515/615), separate latches, or acombination of these. Once deployed, it can begin to provide power tothe spacecraft, including being rotated or angled by the solar arraydrive assembly.

As noted above, in some embodiments the hinges 515/615 can incorporate amotor that can deploy and also re-stow (and subsequently re-deploy) thesolar array. The ability to re-stow a solar array can be useful when aspacecraft is repositioned and can also be useful in other applications,such when a lunar rover or similar vehicle using such an array ismoving. The following section considers embodiments with motorizedhinges in more detail.

FIG. 12 is a side view of an embodiment for a motorized large angleflexure pivot hinge 1215 that can be used in the frame structure for theZ-fold flexible blanket solar array structures presented here. Themotorized flexure hinge 1215 includes a pair of lugs or legs 1271 and1273 that can mount directly to corresponding frame members and that areeach mounted to a central pivot portion, with the leg 1271 attached tothe outer rim 1267 and the leg 1273 attached to an inner rim that isbehind the outer rim 1267 (shown as inner rim 1287 in the view of FIG.13 ). The outer rim 1267 is attached to an outer center hub 1269 atcenter by “leaf spring” spokes 1265 or other flexure elements that flexas the hinge pivots. Alternating with the longer flexures 1265 areshorter “leaf spring” flexures 1264 that are connected to the outercenter hub, but, rather than extend all the way to the outer rim 1267,the flexures 1264 are shorter and attach at their outer ends to atriangular piece 1267. The motorized large angle flexure pivot hinge1215 can also incorporate a retractable latch structure that, in theshown embodiment, includes a bimetallic latch arm 1263 that is mountedon the outer rim 1267 and extends into the page toward the inner rim1287 (FIG. 13 ) and is adjacent to a notch 1261 in the outer rim 1267.The inner part of the motorized flexure hinge can be similarlystructured, aside from the bimetallic latch arm 1263. The motorizedflexure hinge 1215 structures are mounted in pairs coaxially throughshorter half of the “leaf spring” flexures 1264 at the distal end, nearthe rim 1267, so that the triangular structure is between inner andouter parts and attached to the shorter flexures 1264 to either side.(More detail on an embodiment for the flexure pivot, but not includingthe retractable latch structure such as the bimetallic latch arm 1263arrangement, is Spanoudakis et al., “Large Angle Flexure PivotDevelopment for Future Science Payloads”, in the Proceedings of the18^(th) European Space Mechanisms and Tribiology Symposium 2019, whichis incorporated by reference herein.)

The rim notch 1261 and bimetallic latch arm 1263 are offset radiallyalong the outer rim 1267 by an angle φ, where, depending on theembodiment, this could be in the range of 0°-20° (e.g., 8°). Theembodiment of FIG. 12 has two rim notch 1261/bimetallic latch arm 1263pairs offset by 180°. When the motorized flexure hinge 1215 pivots, theouter rim 1267 and leg 1271 rotate clockwise and the inner rim 1287 andleg 1273 rotates counter-clockwise relative to each other. Once the rimshave undergone relative rotations of ±(90°−φ), the bimetallic latch arms1263 on the outer rim 1267 will align with the notches on the inner. Thebimetallic arms can passively latch the motorized flexure hinge 1215 byuse of stored strain energy into a deployed position with a relativeangle of (180°−2φ) between the legs 1271 and 1273. This can beillustrated with reference to FIG. 13 .

The motorized flexure hinges 1215 can use double gear reduction motorsto provide high torque that can be used to tension the solar arrayblanket. The motorized flexure hinges 1215 hinge can be used not only tostow and deploy the array, but also on the deployment stroke to providea tensioning force, reducing or eliminating the need of a separaterelease and tensioning mechanism, saving on costs, mass, and integrationtime and reducing risk due to potential lack of function on-orbit. Toprovide and assist with tensioning in such an arrangement, the solararray blanket can have a degree of elasticity allowing it to stretch inthe direction of the deployment. The frame structure can be engineeredso that its deployed length would be somewhat longer (such as a fewinches) that the solar array blanket's relaxed length. The blanket wouldbe configured to have elasticity such that it would be sufficientlycompliant itself, include a string of compliance (a bungee cord type ofstructure), or a combination of these such that when stretched the solararray blanket can achieve the desired blanket tension, resulting in thecompliant spring displacement to ensure tension (such as a few inches)is maintained throughout the orbit with changing temperatures on thestructure and blanket.

FIG. 13 is an oblique view of the motorized flexure hinge 1215 showingthe features of FIG. 12 and also the inner rim 1287, but without legs1271 and 1273 attaching the rims to the frame structure. Legs 1271 and1273 aside, the features of FIG. 12 are repeated for the outer section,but with only the rim 1267 and bimetallic latch arm 1263 explicitlynumbered. The motor for rotating the outer rim 1267 relative to theinner rim 1287 is within the structure and not visible in the figures.For the inner section, a portion of the inner rim 1287 is visible inFIG. 13 and includes one of its notches 1281. FIG. 13 shows thestructure in an intermediate state between stowed and deployed when theouter rim 1267 has rotated relative to the inner rim 1287 by around 30°in the direction of the arrow. The bimetallic latch arm 1263 is,depending on the embodiment, actively raised from or gliding along theouter surface of the inner rim 1287. Once the inner rim 1287 and outerrim 1267 have rotated through a relative angle of (180°−2φ), thebimetallic latch arm 1263 will be over, and can be lower or snap downinto the notch 1281 to engage the latch and notch to latch the framestructure into its deployed position. A similar bimetallic latch arm1263 and notch 1281 on the other side can similarly engage.

The motorized flexure hinge 1215 can be formed in a number of ways,depending on the embodiment. For example, an additively manufacturedparts can be manufactured through use of Electrical Discharge Machining(EDM) to create the hinges. In other embodiments, a 3D printed hinge canbe manufactured from a reinforced nylon or many other materials as oneassembly, including the lugs or legs 1271 and 1273, which can be bondeddirectly into the frame member tubes to minimize parts and fasteners.

Although in the of embodiments of FIGS. 12 and 13 the element 1263 hasbeen referred to as a bimetallic latch arm 1263, in other embodiments itcould just be a passively flexed member that would latch into place oncethe notch 1281 is aligned with the member 1263. However, the use of abimetallic latch arm is one embodiment of a mechanism that allows forthe latch to be released, allowing the deployed structure to be releasedand folded back up. This is illustrated with respect to FIGS. 14A and14B.

FIGS. 14A and 14B respectively illustrate a bimetallic latch arm 1263 ina released state and a latched state. In the embodiments presented here,the latched state of FIG. 14B corresponds to the relaxed state where thebimetallic latch arm 1263 is at or relatively near the ambienttemperature and FIG. 14A corresponds to the active state when the arm isarced into a released configuration, but alternate embodiments couldreverse these.

The bimetallic latch arm 1263 is formed of a pair of metal strips 1263 aand 1263 b with differing coefficients of heat expansion. Because ofthis, depending on the temperature of the bimetallic latch arm 1263, itscurvature will vary. On the one end, the bimetallic latch arm 1263 isattached to the outer rim 1267 by a bolt, rivet, or similar attachmentdevice 1291 so that it is held fixed on the one end, but allowed to flexalong the length of the bimetallic latch arm 1263. The attachment 1291is made of a material that can conduct heat to the pair of metal strips1263 a and 1263 b and can include a heating element 1293. When at ornear the ambient temperature, both of metal strips 1263 a and 1263 b areof the same or near the same length, so the bimetallic latch arm 1263 isstraight or near straight so that it extends from the outer rim 1267into the notch of the inner rim 1287 to latch the structure into thedeployed configuration of FIG. 14B. In alternate embodiments, thebimetallic latch arm 1263 could be attached to the inner rim 1287 andlatch into the outer rim 1267.

When the heating element 1293 is activated, such as by applying acurrent, to heat the bimetallic latch arm 1263 to, for example, in arange of 225-500 C, the bimetallic latch arm 1263 will arch upward asillustrated in FIG. 14A so that the Z-fold flexible blanket solar arrayframe can be re-stowed by the motorized flexure hinge 1215, folding thestructure back into a configuration such as shown in FIG. 5A or 6A. Oncethe motorized flexure hinge 1215 rotates the outer rim 1267 relative theinner rim 1287 enough so that the bimetallic latch arm 1263 is no longeris aligned with the notch, the bimetallic latch arm 1263 can either bemaintained in an arched configuration or allowed to relax and rest uponthe inner rim, where this is a choice between the energy to maintain thebimetallic latch arm 1263 in the flexed state and energy needed toovercome the friction of the bimetallic latch arm 1263 on the inner rim1287. As the inner rim may of a material (e.g., a nylon material) ofrelatively low drag coefficient, the bimetallic latch arm 1263 can oftenjust be relaxed and the motorized flexure hinge 1215 be allowed toreturn to its stowed position. Similarly, when deploying the structure,the motorized flexure hinge 1215 can glide along the inner rim 1287until it latches in the notch 1281 or be flexed until these align andthen be relaxed.

The use of a flexible bimetallic strip latch arm 1263 that passivelylatches upon deployment by falling into a groove or notch 1281 like akeyway, and that lifts out of the groove when current is applied to aheating element 1293 affixed to the bimetallic latch arm 1263, can be aneffective embodiment for a retractable latch structure. The samecontroller as the motor control board used to power and control themotorized flexure hinges 1215 can also be used to apply current to theheating element 1293. The bimetallic latch arm 1263 can be designed tonot deflect out of the locking notch 1281 through temperature variationsexperienced on-orbit and only release when a current is applied to theheating element 1293 so that it has a sufficiently greater temperaturethan would be experienced on-orbit. The deflected bimetallic strip latcharm 1263 of FIG. 14A could raise high enough to land on a ramp, in someembodiments, which would fully un-latch upon rotating the hinge with themotor.

As discussed above, as the notch 1281 and arm bimetallic latch arm 1263can be offset by radial angle φ on the outer rim and inner rim 1287,when latched into place the yokes and intermediate members will not beflat, but rather at an obtuse angle of (180°−2φ). As the solar arrayblanket or blankets are only connected at the ends to the yokestructures, but not the intermediate frame members, the solar arrayblanker or blankets can still be flat, and the frame used to tension theblanket or blankets. This can be illustrated with respect to FIGS. 15and 16 .

FIGS. 15 and 16 illustrate an embodiment for a deployed Z-fold flexibleblanket solar array structure in which the yokes and intermediatemembers are at less than 180° when latched in a deployed configuration.FIG. 15 again shows the Z-fold flexible blanket solar array structureattached to a spacecraft 10, but it could also be mounted to other hoststo provide power, such as to a lunar rover or other vehicle. The inboardT-yoke 1511 is attached to the mount 1551, with one or more (three inthis example) intermediate members 1513 connected between the inboardT-yoke 1511 and the outboard T-yoke 1517. In this embodiment the inboardT-yoke 1511, intermediate members 1513, and outboard T-yoke 1517 membersof the frame are connected by motorized flexure hinges 1515, such asdescribed with respect to FIGS. 12, 13, 14A, and 14B. In alternateembodiments, non-motorized hinges as described with respect to earlierembodiments can be used for the hinges and can, for example, include aspring to deploy the structure, but the use of motorized hinges andretractable latches allow for the structure to be re-stowed. Althoughnot noted in FIG. 15 , the structure can again include solar arraysmounted on one or both of the inboard T-yoke 1511 and the outboardT-yoke 1517.

One or more blanket solar arrays 1521 are stretched between the inboardT-yoke 1511, where one end is attached at 1591, and outboard T-yoke1517, where the other end is attached at 1593, but where the one or moreblanket solar arrays 1521 are not attached to the intermediate members1513. Relative to FIG. 5D, rather the different frame members connectingat 180° when deployed, they are at an angle φ relative to the horizontal(as represented in FIG. 15 ) so that the angle between the members isthe obtuse angle of (180°−2φ) where, in this embodiment, φ could be inthe range of 0°-20° (e.g., 8°). When viewed from above, the deployedstructure of FIG. 15 would appear as in FIG. 5C for a two blanketembodiment, with a blanket to either side of the central frame member,or as in FIG. 6B for a single blanket embodiment. When stowed, these twoembodiments would be similar to FIG. 5A or 6A and held in place byhold-down 1519 and could again incorporate folding arms as illustratedwith respect to FIGS. 9A-9D. When at an intermediate state whileextending from the stowed configuration to the deployed configuration(or being folded up to be re-stowed), the structure of FIG. 15 would besimilar to the representation of FIG. 5B, but where the deployment wouldlatch into place in the configuration of FIG. 15 rather than a full180°.

As noted above, the solar array blanket 1521 can be configured to have adegree of elasticity, allowing it to stretch by, perhaps, a few inchesin order to provide tensioning. By not deploying frame members to a full180°, the motorized flexure hinges 1515 can use their motor torque totranslate the extension into blanket tension. The amount of torque fromthe motorized flexure hinges 1515 will typically decrease in a somewhatsinusoidal manner near 180°, so that it in some embodiments it can bepreferable for the frame to be extended somewhat less than 180°.

The embodiment of FIG. 15 , where the frame members (inboard T-yoke1511, intermediate members 1513, and outboard T-yoke 1517) arerelatively inclined at an angle of less than 180° (e.g., obtuse anglesbetween 170° and 160°), can hold the deployed solar array blanket orblankets 1521 stretched taut between the mounting points 1591 and 1593.The frame members in this configuration can have the ability to flex tosome degree, so that solar array blanket or blankets 1521 can be keptstretched taut as they expand or contract based on conditions such asthe degree of exposure to the sun.

The frame members (inboard T-yoke 1511, intermediate members 1513, andoutboard T-yoke 1517) of FIG. 15 can again formed of hollow graphiterectangular tubes with dimensions of a few inches and a wall thicknessof 10s of mils or of other beam shapes (e.g., I-beam, round or ovalcross-sections). For embodiments that require additional rigidity, suchfor larger arrays, the frame members can use a bean architecture inwhich the beams are wider in the horizontal directions, as illustrate inFIG. 16 .

FIG. 16 is a top view of the embodiment of FIG. 15 when inboard T-yoke1511, intermediate members 1513, and outboard T-yoke 1517 use arectangular beam structure, but with the solar blanket or blankets andany yoke mounted solar arrays not shown. Each of the frame members inthis embodiment have a rectangular cross-section that wider in thehorizontal directions of FIG. 16 (i.e., the plane of the deployed solararray blanket) then in the direction into the page (i.e., the side viewof FIG. 15 ). Each of the individual beams can again be formed of hollowgraphite rectangular tubes and have a cross-section with an aspect ratioof 3-4 to 1, for examples, such 6-10 inches wide and −1-2 inches thick,with the sizes selected for a given embodiment depend on therequirements of the application. The frame members are attached to oneanother and to the spacecraft, vehicle, or other device by a pair ofcoaxially mounted motorized flexure hinges 1515, where one of the pairare attached to one of the beam pairs on either side and are connectedto each other an axle. Alternate embodiments can use other hingestructures and the dual beam frame member structure can be used for the“straight” embodiments such as described with respect to FIGS. 5A-5D orFIG. 6A-6B. In other alternate embodiments, the structures describedwith respect to FIGS. 15 and 16 can also incorporate fold down arms asin FIGS. 9A-9C.

The embodiments described with respect to FIGS. 12-16 can have a numberfurther advantages over prior art arrangements, including the ability tore-stow the array, a property that can be useful, for example, to thelunar rover or similar vehicle that has the requirement to be drivenaround, where the array can be deployed to charge up the host and thenre-stowed to drive around again. By having a deployed state in which theframe members are not flat (i.e., not 180°), the frame structure ofFIGS. 15 and 16 provide the ability to tension the solar array blanketwithout need of a separate mechanism made up of release devices andtensioning devices. The structure can leverage the high torque of motorsof the motorized flexure hinges 1515 to generate tensioning force in thesolar array blanket, eliminating the two separate passive mechanisms ofa release device and tensioning springs. By having the frame structureable to deploy further than the solar array blanket's length, by a fewinches for example, and providing the solar array blankets to include a“lower-stiffness” spring, the force generated by the differential lengthcan help to ensure through various parts of an orbit and associatedtemperatures changes that the frame structure maintains tension on thesolar array blanket.

The rigidity of the deployed Z-fold flexible blanket solar arraystructure can also provide the gimbal control electronics 218 with theability to adjust the solar array angle as attached at a solar arraydrive assembly within the mount 1551 to allow for roll-steering thespace craft for coverage or fine tuning the solar array orientation forvarious parts of the orbit to maximize the cell efficiency. Suchroll-steering can be useful for the spacecraft with changing coverageareas. The rigidity of the frame structure can similarly provide theability to adjust for solar pressure differentials between the wings ofa two-array embodiment (as at 265 of FIGS. 3 and 4 ) to reduce the needfor expending valuable propellants for station keeping.

The control circuitry for the Z-fold flexible blanket solar arraystructure, for one or both of the motorized flexure hinges 1515 and theheating element 1293 of the metal strips 1263 a and 1263 b can be partof the spacecraft's (or other host's) control circuit; be on the solararray structure; or a combination of these. Moving the control circuitryto the frame structure can have a number of advantages as this can allowfor the individual control signals to be generated on the arraystructure, reducing the amount of wiring and also simplifying theinteraction with the spacecraft so that individual spacecrafts controlcircuitry need not account for differences in array structures (i.e.,differing numbers or types of motorized flexure hinges 1515 orretractable latch structures). FIG. 17 illustrates one such embodiment.

FIG. 17 illustrates an embodiment for the locating of the controller andwiring for the motorized flexure hinges and releasable latches. Thecontrol circuitry for the array controller 1701 is located on the frameof the Z-fold flexible blanket solar array structure. In the embodimentof FIG. 17 , the system controller 1701 is mounted on the inner yoke1511, but other placements can be used. The system controller 1701 isconnected to a wiring harness or harnesses 1703 that can be used toprovide control signals to the motorized flexure hinges 1515 includingretractable latch structures through a minimal number of slip rings inthe solar array drive assembly (SADA) in the mount 1551. The wiringharness also connects the on-array system controller to the controlcircuitry of the spacecraft, lunar rover, or other host on which thesolar array structure is mounted. The communication between the host tothe Z-fold flexible blanket solar array structure's on-array systemcontroller 1701 can be as simple as signals to deploy and re-stow, alongwith confirmations, with the system controller 1701 converting these tothe appropriate number and type of signals to supply to the motorizedflexure hinges 1515 and heating element 1293 or other retractable latchstructure release mechanism.

The Z-fold flexible blanket solar array structure embodiments of FIGS.12-17 can operate as described above with respect to FIGS. 10 and 11 forthe stowing and deploying of the solar array. Additionally, followingdeployment at steps 1113 and 1115, the structure can be restowed andsubsequently redeployed, as described with respect to FIG. 18 .

FIG. 18 is a flowchart of an embodiment for a method of operating aretractable Z-fold flexible blanket solar array structure. FIG. 18 picksup the flow at steps 1813 and 1815, which can correspond to step 1113and 1115 of FIG. 11 . If the flow of FIG. 18 is the initial deployment,or the hold-downs have been reattached, the hold-downs are releasedprior to step 1813. Additionally, if the embodiment includes foldablearms, these can have been deployed (or redeployed) prior to step 1813 inan initial deployment phase.

At step 1813, in response to control signals from on-array controller1701, which are in turn in response to a deploy signal from thespacecraft 10 or other host, the motorized flexure hinges 1515 areenabled and the Z-fold flexible blanket solar array structure expandedout to its deployed configuration. If the latches are set when in thestowed configuration, the latches are also released at step 1813. Oncethe frame has expanded into the deployed configuration, it is latched inplace as step 1815, such as by the bimetallic strip 1263 slotting in thenotch 1281 on the rims of the motorized flexure hinge 1215. In thedeployed configuration, the solar array blanket can then provide powerto the spacecraft 10 or other host.

At some later time, the spacecraft, lunar rover, or other apparatususing the solar array structure issues a stow command to the Z-foldflexible blanket solar array structure at step 1817. The re-stow commandcan be received at the on-array system controller 1701 that can issuerelease commands to the latches and then activate the motorized flexurehinge 1215. At step 1819 the retractable latch structures are release,such as by applying a current to the heating element 1293 of thebimetallic strip 1263 to heat the bimetallic strip 1263, allowing it torelease. Once the latches are released, the motorized flexure hinge 1215fold up the Z-fold flexible blanket solar array structure to re-stow itat step 1811. Depending on the embodiment, the bimetallic strips 1263can be maintained in a flexed state while the Z-fold flexible blanketsolar array structure is folded up or they can be relaxed once themotorized flexure hinges 1215 rotate sufficiently so that the bimetallicstrips 1263 are clear of the notches 1281. Depending on the embodiment,once folded up the solar array structure can be reattached to thehold-downs 1519; held in place by additional notches like 1281, butlocated to hold the structure in the stowed configuration; held in placeby the motorized flexure hinge 1215; or a combination of these. If thestructure also includes foldable arms, such as illustrated in FIGS.9A-9D, these can also either be left extended or also folded up in step1821.

Once the Z-fold flexible blanket solar array structure is re-stowed, thespacecraft, lunar rover, or other hosting apparatus can be relocated atstep 1823 without worrying about the solar array structure being damagedas easily or inhibiting movement. Once the movement is finished or poweris need, the Z-fold flexible blanket solar array structure can beredeployed by sending a command to the on-array structure controller1701 at step 1825 and the flow can loop back to step 1813 fordeployment.

One embodiment includes an apparatus comprising a foldable solar arrayfoldable in a first direction in an accordion manner to form a stowedpack having a plurality of folds and a deployable frame structureconnected to the foldable solar array. The deployable frame structureincludes: a T-shaped yoke structure having a central member and across-member arm, the cross-member arm attached to the foldable solararray at a first end in the first direction and the central memberconfigured to mount to a spacecraft; a T-shaped end structure having acentral member and a cross-member arm, the cross-member arm attached tothe foldable solar array at a second end in the first direction; one ormore rigid beams; and a plurality of hinges connecting the T-shaped yokestructure, the one or more rigid beams, and the T-shaped end structure,including a first hinge connected to the T-shaped yoke structure and asecond hinge connected to an end central member of the T-shaped endstructure. The plurality of hinges are configured to: fold thedeployable frame structure into a stowed configuration in which thestowed pack is stored between the cross-member arm of the T-shaped yokestructure and the cross-member arm of the T-shaped end structure; andextend the deployable frame structure in the first direction into adeployed configuration in which the foldable solar array is tensionedbetween the cross-member arm of the T-shaped yoke structure and thecross-member arm of the T-shaped end structure.

One embodiment includes an apparatus including a foldable solar arrayfoldable in a first direction in an accordion manner to form a stowedpack having a plurality of folds and a deployable frame structureconnected to the foldable solar array. The deployable includes: aT-shaped yoke structure having a cross-member arm attached to thefoldable solar array at a first end in the first direction, the T-shapedyoke structure configured to mount to a spacecraft; and a T-shaped endstructure having a cross-member arm attached to the foldable solar arrayat a second end in the first direction, the T-shaped end structurehaving a first solar array on a surface of the cross-member arm. Thedeployable frame structure is configured to: extend in the firstdirection into a deployed configuration in which the foldable solararray tensioned between the cross-member arm of the T-shaped yokestructure and the cross-member arm of the T-shaped end structure; andfold into a stowed configuration in which the stowed pack is storedbetween the cross-member arm of the T-shaped yoke structure and thecross-member arm of the T-shaped end structure and in which the firstsolar array faces outward from a surface of a spacecraft to which theapparatus is attached.

One embodiment includes a method comprising: rotating outward first andsecond arms of a yoke structure to form a T-shaped yoke structure of astowed solar array; and rotating outward first and second arms of an endstructure to form a T-shaped end structure of the stowed solar array.The stowed solar array includes: a first foldable solar array folded ina first direction in an accordion manner into a first pack stowedbetween the first arm of the yoke structure and the first arm of the endstructure, a second foldable solar array folded in the first directionin an accordion manner into a second pack stowed between the second armof the yoke structure and the second arm of the end structure, and adeployable frame structure including the yoke structure and the endstructure. Subsequent to rotating outward the arms of the yoke structureand the arms of the end structure, the method also includes extendingthe deployable frame structure in the first direction into a deployedconfiguration in which the first foldable solar array is tensionedbetween the first arm of the T-shaped yoke structure and the first armof the T-shaped end structure and the second foldable solar array istensioned between the second arm of the T-shaped yoke structure and thesecond arm of the T-shaped end structure.

Further embodiments include an apparatus having one or more foldablesolar arrays; and a deployable frame structure connected to the one ormore solar array blankest and configured to mount to a spacecraft. Thedeployable frame structure includes a solar array and is configured to:hold the one or more foldable solar arrays in a stowed configuration asa corresponding one or more stowed packs; while holding the one or morefoldable solar arrays in the stowed configuration, provide power fromthe solar array to a spacecraft to which the deployable frame structureis mounted; and extend the one or more foldable solar arrays from thestowed configuration into a deployed configuration.

Additional embodiments include an apparatus having a foldable solararray configured to fold in a first direction to form a stowed pack in astowed configuration and to provide power to a host when in a deployedconfiguration and a deployable frame structure connected to the foldablesolar array. The deployable frame structure includes: a yoke structureattached to the foldable solar array at a first end in the firstdirection and configured to mount to the host; an end structure attachedto the foldable solar array at a second end in the first direction; andone or more intermediate frame member. A plurality of motorized hingesconnect the yoke structure, the one or more intermediate frame members,and the end structure, each of the motorized hinges including aretractable latch structure, the plurality of motorized hingesconfigured to: extend the deployable frame structure in the firstdirection from the stowed configuration into the deployed configuration,in which the foldable solar array is tensioned between the yokestructure and the end structure, and latch the deployed frame structureinto the deployed configuration with the retractable latch structures;and release the retractable latch structures and re-fold the deployableframe structure into the stowed configuration.

Embodiments also include a method that comprises receiving, at a solararray structure in a stowed configuration, a deployment command. Thesolar array structure includes: a foldable solar array configured tofold in a first direction to form a stowed pack in the stowedconfiguration and to provide power to a host when in a deployedconfiguration; a deployable frame structure connected to the foldablesolar array, including a yoke structure attached to the foldable solararray at a first end in the first direction and configured to mount tothe host, an end structure attached to the foldable solar array at asecond end in the first direction, and one or more intermediate framemembers; and a plurality of motorized hinges connecting the yokestructure, the one or more intermediate frame members, and the endstructure, each of the motorized hinges including a retractable latchstructure. In response to the deployment command, the method includes:extending by the motorized hinges of the deployable frame structure inthe first direction from the stowed configuration into the deployedconfiguration, in which the foldable solar array is tensioned betweenthe yoke structure and the end structure; and latching the deployedframe structure into the deployed configuration with the retractablelatch structures. Subsequent to latching the deployed frame structureinto the deployed configuration, the method further includes receiving astow command and, in response to the stow command, releasing theretractable latch structures and, subsequent to releasing theretractable latch structures, re-folding the deployable frame structureinto the stowed configuration.

Further embodiments include an apparatus, comprising: a solar arrayfoldable in a first direction to form a stowed pack and configured toelastically extend in the first direction when in a deployedconfiguration; a deployable frame structure connected to the solararray; and a plurality of hinges. The deployable frame structureincludes: a T-shaped yoke structure having a central member and across-member arm, the cross-member arm attached to the solar array at afirst end in the first direction and the central member configured tomount to a host; a T-shaped end structure having a central member and across-member arm, the cross-member arm attached to the solar array at asecond end in the first direction; and one or more intermediate framemembers. The hinges are configured to: fold the deployable framestructure into a stowed configuration in which the stowed pack is storedbetween the cross-member arm of the T-shaped yoke structure and thecross-member arm of the T-shaped end structure; and extend thedeployable frame structure in the first direction into a deployedconfiguration in which the T-shaped yoke structure, the one or moreintermediate frame members, and the T-shaped end structure are connectedto one another to elastically extend the solar array to be tensionedbetween the cross-member arm of the T-shaped yoke structure and thecross-member arm of the T-shaped end structure.

For purposes of this document, it should be noted that the dimensions ofthe various features depicted in the figures may not necessarily bedrawn to scale.

For purposes of this document, reference in the specification to “anembodiment,” “one embodiment,” “some embodiments,” or “anotherembodiment” may be used to describe different embodiments or the sameembodiment.

For purposes of this document, a connection may be a direct connectionor an indirect connection (e.g., via one or more other parts). In somecases, when an element is referred to as being connected or coupled toanother element, the element may be directly connected to the otherelement or indirectly connected to the other element via interveningelements. When an element is referred to as being directly connected toanother element, then there are no intervening elements between theelement and the other element. Two devices are “in communication” ifthey are directly or indirectly connected so that they can communicateelectronic signals between them.

For purposes of this document, the term “based on” may be read as “basedat least in part on.”

For purposes of this document, without additional context, use ofnumerical terms such as a “first” object, a “second” object, and a“third” object may not imply an ordering of objects, but may instead beused for identification purposes to identify different objects.

For purposes of this document, the term “set” of objects may refer to a“set” of one or more of the objects.

The foregoing detailed description has been presented for purposes ofillustration and description. It is not intended to be exhaustive or tolimit the subject matter claimed herein to the precise form(s)disclosed. Many modifications and variations are possible in light ofthe above teachings. The described embodiments were chosen in order tobest explain the principles of the disclosed technology and itspractical application to thereby enable others skilled in the art tobest utilize the technology in various embodiments and with variousmodifications as are suited to the particular use contemplated. It isintended that the scope of be defined by the claims appended hereto.

What is claimed is:
 1. An apparatus, comprising: a foldable solar arrayconfigured to fold in a first direction to form a stowed pack when in astowed configuration and to provide power to a host when in a deployedconfiguration; a deployable frame structure connected to the foldablesolar array, including: a yoke structure attached to the foldable solararray at a first end in the first direction and configured to mount tothe host; an end structure attached to the foldable solar array at asecond end in the first direction; and one or more intermediate framemembers; and a plurality of motorized hinges connecting the yokestructure, the one or more intermediate frame members, and the endstructure, each of the motorized hinges including a retractable latchstructure, the plurality of motorized hinges configured to: extend thedeployable frame structure in the first direction from the stowedconfiguration into the deployed configuration, in which the foldablesolar array is tensioned between the yoke structure and the endstructure, and latch the deployed frame structure into the deployedconfiguration with the retractable latch structures; and release theretractable latch structures and re-fold the deployable frame structureinto the stowed configuration.
 2. The apparatus of claim 1, wherein eachof the motorized hinges includes: a first rim connected to one of theyoke structure, the end structure, or one of the intermediate framemembers; a second rim connected to another of one of the yoke structure,the end structure, or one of the intermediate frame members; and a motorconfigured to rotate the first rim relative to the second rim.
 3. Theapparatus of claim 2, wherein, for each of the motorized hinges, theretractable latch structure includes: a notch in the second rim; and abimetallic strip attached to the first rim and configured to engage withthe notch to latch the deployed frame structure into the deployedconfiguration.
 4. The apparatus of claim 1, wherein the solar array isconfigured to elastically extend in the first direction when in adeployed configuration
 5. The apparatus of claim 1, wherein: the yokestructure is a T-shaped yoke structure having a central member and across-member arm, the cross-member arm attached to the foldable solararray at a first end in the first direction and the central memberconfigured to mount to the host; and the end structure a T-shaped endstructure having a central member and a cross-member arm, thecross-member arm attached to the foldable solar array at a second end inthe first direction.
 6. The apparatus of claim 5, wherein, when latchedinto the deployed configuration, one or more of the motorized hingesconfigure to latch at an obtuse angle.
 7. The apparatus of claim 1,wherein the host is a spacecraft and the deployable frame structure isattached to the host by a solar array drive assembly configured torotate the deployable frame structure to roll-steer the space craft. 8.The apparatus of claim 1, further comprising: a controller mounted onthe deployable frame structure configured to receive instructions fromthe host and control operation of the motorized hinges and retractablelatch structures in response to the instructions from the host.
 9. Theapparatus of claim 1, wherein the host is a spacecraft.
 10. Theapparatus of claim 1, wherein the host is a vehicle.
 11. A method,comprising: receiving, at a solar array structure in a stowedconfiguration, a deployment command, the solar array structureincluding: a foldable solar array configured to fold in a firstdirection to form a stowed pack in the stowed configuration and toprovide power to a host when in a deployed configuration; a deployableframe structure connected to the foldable solar array, including a yokestructure attached to the foldable solar array at a first end in thefirst direction and configured to mount to the host, an end structureattached to the foldable solar array at a second end in the firstdirection, and one or more intermediate frame members; and a pluralityof motorized hinges connecting the yoke structure, the one or moreintermediate frame members, and the end structure, each of the motorizedhinges including a retractable latch structure; in response to thedeployment command: extending by the motorized hinges of the deployableframe structure in the first direction from the stowed configurationinto the deployed configuration, in which the foldable solar array istensioned between the yoke structure and the end structure; and latchingthe deployed frame structure into the deployed configuration with theretractable latch structures; subsequent to latching the deployed framestructure into the deployed configuration, receiving a stow command; andin response to the stow command: releasing the retractable latchstructures; and subsequent to releasing the retractable latchstructures, re-folding the deployable frame structure into the stowedconfiguration.
 12. An apparatus, comprising: a solar array foldable in afirst direction to form a stowed pack and configured to elasticallyextend in the first direction when in a deployed configuration; adeployable frame structure connected to the solar array, including: aT-shaped yoke structure having a central member and a cross-member arm,the cross-member arm attached to the solar array at a first end in thefirst direction and the central member configured to mount to a host; aT-shaped end structure having a central member and a cross-member arm,the cross-member arm attached to the solar array at a second end in thefirst direction; and one or more intermediate frame members; and aplurality of hinges connecting the T-shaped yoke structure, the one ormore intermediate frame members, and the T-shaped end structure, theplurality of hinges configured to: fold the deployable frame structureinto a stowed configuration in which the stowed pack is stored betweenthe cross-member arm of the T-shaped yoke structure and the cross-memberarm of the T-shaped end structure; and extend the deployable framestructure in the first direction into the deployed configuration inwhich the T-shaped yoke structure, the one or more intermediate framemembers, and the T-shaped end structure are connected to one another toelastically extend the solar array to be tensioned between thecross-member arm of the T-shaped yoke structure and the cross-member armof the T-shaped end structure.
 13. The apparatus of claim 12, whereinone or more of the hinges are motorized hinges configured to: deploy theframe structure from the stowed configuration to the deployedconfiguration; and re-fold the frame structure from the deployedconfiguration to the stowed configuration.
 14. The apparatus of claim13, wherein each of the motorized hinges includes a retractable latchstructure configured to latch the frame structure into the deployedconfiguration.
 15. The apparatus of claim 14, wherein the motorizedhinges include a retractable latch structure configured to latch thedeployable frame structure into a deployed configuration.
 16. Theapparatus of claim 15, wherein each of the motorized hinges comprise: afirst rim connected to one of the yoke structure, the end structure, orone of the intermediate frame members; a second rim connected to anotherof one of the yoke structure, the end structure, or one of theintermediate frame members; and a motor configured to rotate the firstrim relative to the second rim.
 17. The apparatus of claim 16, wherein,for each of the motorized hinges, the retractable latch structureincludes: a notch in the second rim; and a bimetallic strip attached tothe first rim and configured to engage with the notch to latch thedeployed frame structure into the deployed configuration.
 18. Theapparatus of claim 12, wherein when the frame structure is extended inthe first direction into the deployed configuration the T-shaped yokestructure, the one or more intermediate frame members, and the T-shapedend structure are connected to one another at obtuse angles.
 19. Theapparatus of claim 12, wherein the host is a spacecraft.
 20. Theapparatus of claim 12, wherein the host is a vehicle.